On the fly laser shock peening

ABSTRACT

An on the fly method of laser shock peening a gas turbine engine part by continuously moving a metallic gas turbine engine part while continuously firing a stationary laser beam, which repeatably pulses between relatively constant periods, on a portion of the part with sufficient power to vaporize material on the surface of the portion of the part with the pulses around laser beam spots formed by the laser beam on the surface and form a region having deep compressive residual stresses extending into the part from the laser shock peened surface. Flowing a curtain of water over the surface upon which the laser beam is firing while moving the part until the laser shock peened surface is completely covered by laser beam spots at least once. The surface may covered by a paint which is then the material used to produce the plasma or the surface may be unpainted and the metal of the part is used to produce the plasma. The part such a fan or compressor blade may be moved linearly to produce at least one row of overlapping circular laser beam spots having generally equally spaced apart linearly aligned center points.

The present Application deals with related subject matter in U.S. patentapplication Ser. No. 08/319,346, entitled “LASER SHOCK PEENED ROTORCOMPONENTS FOR TURBOMACHINERY”, filed Oct. 6, 1994, assigned to thepresent Assignee, now U.S. Pat. No. 5,492,447.

The present Application deals with related subject matter in U.S. patentapplication Ser. No. 08/373,133, now U.S. Pat. No. 5,591,009 entitled“LASER SHOCK PEENED GAS TURBINE ENGINE FAN BLADE EDGES”, filed Dec. 16,1994, assigned to the present Assignee.

BACKGROUND OF THE INVENTION

1. Field of the Invention

This invention relates to laser shock peening of gas turbine engineparts and, more particularly, for airfoil leading and trailing edgessuch as on fan and compressor blades having localized compressiveresidual stresses imparted by laser shock peening.

2. Description of Related Art

Gas turbine engines and, in particular, aircraft gas turbine enginesrotors operate at high rotational speeds that produce high tensile andvibratory stress fields within the blade and make the fan bladessusceptible to foreign object damage (FOD). Vibrations may also becaused by vane wakes and inlet pressure distortions as well as otheraerodynamic phenomena. This FOD causes nicks and tears and hence stressconcentrations in leading and trailing edges of fan blade airfoils.These nicks and tears become the source of high stress concentrations orstress risers and severely limit the life of these blades due to HighCycle Fatigue (HCF) from vibratory stresses. FOD damage may also resultin a loss of engine due to a release of a failed blade. It is alsoexpensive to refurbish and/or replace fan blades and, therefore, anymeans to enhance the rotor capability and, in particular, to extendaircraft engine fan blade life is very desirable. The present solutionto the problem of extending the life of fan blades is to design adequatemargins by reducing stress levels to account for stress concentrationmargins on the airfoil edges. This is typically done by increasingthicknesses locally along the airfoil leading edge which adds unwantedweight to the fan blade and adversely affects its aerodynamicperformance. Another method is to manage the dynamics of the blade byusing blade dampers. Dampers are expensive and may not protect bladesfrom very severe FOD. These designs are expensive and obviously reducecustomer satisfaction.

Therefore, it is highly desirable to design and construct longer lastingfan and compressor blades, as well as other gas turbine engine parts,that are better able to resist both low and high cycle fatigue thanpresent day parts. The above referenced U.S. patent applications Ser.No. 08/319,346, entitled “LASER SHOCK PEENED ROTOR COMPONENTS FORTURBOMACHINERY”, and U.S. patent application Ser. No. 08/373,133,entitled “LASER SHOCK PEENED GAS TURBINE ENGINE FAN BLADE EDGES”, aredirected towards this end. The latter, more particularly, provides anairfoil of a fan blade with regions of deep compressive residualstresses imparted by laser shock peening at least a radially extendingportion of leading and/or trailing edge surfaces of the fan blade.

The region of deep compressive residual stresses imparted by laser shockpeening of the present invention is not to be confused with a surfacelayer zone of a work piece that contains locally bounded compressiveresidual stresses that are induced by a hardening operation using alaser beam to locally heat and thereby harden the work piece such asthat which is disclosed in U.S. Pat. No. 5,235,838, entitled “Method andapparatus for truing or straightening out of true work pieces”. Thepresent invention uses multiple radiation pulses from high power pulsedlasers to produce shock waves on the surface of a work piece similar tomethods disclosed in U.S. Pat. No. 3,850,698, entitled “AlteringMaterial Properties”; U.S. Pat. No. 4,401,477, entitled “Laser shockprocessing”; and U.S. Pat. No. 5,131,957, entitled “MaterialProperties”. Laser peening as understood in the art and as used herein,means utilizing a laser beam from a laser beam source to produce astrong localized compressive force on a portion of a surface. Laserpeening has been utilized to create a compressively stressed protectionlayer at the outer surface of a workpiece which is known to considerablyincrease the resistance of the workpiece to fatigue failure as disclosedin U.S. Pat. No. 4,937,421, entitled “Laser Peening System and Method”.One issue is manufacturing costs of the laser shock peening processwhich can be prohibitively expensive. The “on the fly” laser shockpeening process of the present invention is designed to provide costsaving methods for laser shock peening.

SUMMARY OF THE INVENTION

A fly method of laser shock peening a gas turbine engine part bycontinuously moving a metallic gas turbine engine part whilecontinuously firing a stationary laser beam, which repeatably pulsesbetween relatively constant periods, on a portion of the part. The laserbeam with sufficient power to vaporize material on the surface of theportion of the part with the pulses around laser beam spots formed bythe laser beam on the surface and to form a region having deepcompressive residual stresses imparted by the laser shock peeningprocess extending into the part from the laser shock peened surface. Themethod further includes flowing a curtain of water over the surface uponwhich the laser beam is firing while moving the part until the lasershock peened surface is completely covered by laser beam spots at leastonce. The surface may be covered by a paint which is then the materialused to produce the plasma or the surface may be unpainted and the metalof the part is used to produce the plasma. The part may be movedlinearly to produce at least one row of overlapping circular laser beamspots having generally equally spaced apart linearly aligned centerpoints and the part may be moved and the laser beam fired to producemore than one row of overlapping circular laser beam spots havinggenerally equally spaced apart linearly aligned center points whereinadjacent rows of spots overlap. The laser beam may be fired and the partmoved so that the center points of adjacent spots in adjacent rows arealso offset from each other a generally equal amount in a directionalong a line on which the center points are linearly aligned.

In another embodiment of the present invention, the painted laser shockpeened surface is laser shock peened using a set of sequences in whicheach sequence the surface is painted and then the part is continuouslymoved while continuously firing a stationary laser beam on the surfacesuch that adjacent laser shock peened circular spots are hit indifferent ones of the sequences in the set. In a more particularembodiment, the laser beam is fired and the part moved so that thecenter points of adjacent spots in adjacent rows are offset from eachother a generally equal amount in a direction along a line on which thecenter points are linearly aligned. A more particular embodiment, eachspot is hit more than one time using more than one set of the sequences.

Another embodiment of the present invention includes a further a step ofremoving remelt formed by the laser firing upon the unpainted lasershock peened surface after it is laser shock peened.

A more particular embodiment of the present invention uses the methodsabove on a gas turbine engine blade such as on the airfoil of a vane orblade of a fan or compressor section having an airfoil with a leadingedge and a trailing edge. The invention may be used along a portion ofthe edge or along the entire edge of the airfoil such that the lasershock peened surface is at least a part of one of the edges extendingradially along at least a portion of and chordwise from the edge.

ADVANTAGES

Among the advantages provided by the present invention is a costefficient method to laser shock peen surfaces of portions of gas turbineengine parts and in particular blades designed to operate in hightensile and vibratory stress fields which can better withstand fatiguefailure due to nicks and tears in the leading and trailing edges of thefan blade and have an increased life over conventionally constructed fanblades. Another advantage of the present invention is that fan andcompressor blades can be constructed with cost efficient methods toprovide commercially acceptable life spans without increasingthicknesses along the leading and trailing edges, as is conventionallydone. The present invention can be advantageously used to refurbishexisting fan and compressor blades with a low cost method for providingsafe and reliable operation of older gas turbine engine fan blades whileavoiding expensive redesign efforts or frequent replacement of suspectfan blades as is now often done or required.

BRIEF DESCRIPTION OF THE DRAWINGS

The foregoing aspects and other features of the invention are explainedin the following description, taken in connection with the accompanyingdrawings where:

FIG. 1 is a cross-sectional illustrative view of an exemplary aircraftgas turbine engine in accordance with the present invention.

FIG. 2 is a perspective illustrative view of an exemplary aircraft gasturbine engine fan blade in accordance with the present invention.

FIG. 2A is a perspective illustrative view of an alternative aircraftgas turbine engine fan blade including a laser shock peened radiallyextending portion of the leading edge in accordance with the presentinvention.

FIG. 3 is a cross sectional view through the fan blade taken along line3—3 as illustrated in FIG. 2.

FIG. 4 is a schematic illustration of a pattern of laser shocked peenedcircular spots on a laser shock peened surface along a leading edge ofthe fan blade in FIG. 2.

FIG. 5 is a schematic illustration of a particular pattern having foursequences of laser shocked peened circular spots that don't overlapwithin a given sequence.

FIG. 6 is a schematical perspective view of the blade of FIG. 1 paintedand mounted in a laser shock peening system illustrated the method ofthe present invention.

FIG. 7 is a partial cross-sectional and a partial schematic view of thesetup in FIG. 7.

DETAILED DESCRIPTION OF THE INVENTION

Illustrated in FIG. 1 is a schematic representation of an exemplaryaircraft turbofan gas turbine engine 10 including a fan blade 8 inaccordance with one embodiment of the present invention. The gas turbineengine 10 is circumferentially disposed about an engine centerline 11and has, in serial flow relationship, a fan section 12, a high pressurecompressor 16, a combustion section 18, a high pressure turbine 20, anda low pressure turbine 22. The combustion section 18, high pressureturbine 20, and low pressure turbine 22 are often referred to as the hotsection of the engine 10. A high pressure rotor shaft 24 connects, indriving relationship, the high pressure turbine 20 to the high pressurecompressor 16 and a low pressure rotor shaft 26 drivingly connects thelow pressure turbine 22 to the fan section 12. Fuel is burned in thecombustion section 18 producing a very hot gas flow 28 which is directedthrough the high pressure and low pressure turbines 20 and 22respectively to power the engine 10. A portion of the air passingthrough the fan section 12 is bypassed around the high pressurecompressor 16 and the hot section through a bypass duct 30 having anentrance or splitter 32 between the fan section 12 and the high pressurecompressor 16. Many engines have a low pressure compressor (not shown)mounted on the low pressure rotor shaft 26 between the splitter 32 andthe high pressure compressor 16. The fan section 12 is a multi-stage fansection, as are many gas turbine engines, illustrated by first, second,and third fan stages; 12 a, 12 b, and 12 c respectively. The fan blade 8of the present invention is designed to be used in a single stage fansection or in any stage of a multi-stage fan section.

Referring to FIGS. 2 and 3, the fan blade 8 includes an airfoil 34extending radially outward from a blade platform 36 to a blade tip 38.The fan blade 8 includes a root section 40 extending radially inwardfrom the platform 36 to a radially inward end 37 of the root section 40.At the radially inward end 37 of the root section 40 is a blade root 42which is connected to the platform 36 by a blade shank 44. The airfoil34 extends in the chordwise direction between a leading edge LE and atrailing edge TE of the airfoil. A chord C of the airfoil 34 is the linebetween the leading LE and trailing edge TE at each cross section of theblade as illustrated in FIG. 3. A pressure side 46 of the airfoil 34faces in the general direction of rotation as indicated by the arrow anda suction side 48 is on the other side of the airfoil and a mean-line MLis generally disposed midway between the two faces in the chordwisedirection.

Referring again to FIG. 2, fan blade 8 has a leading edge section 50that extends along the leading edge LE of the airfoil 34 from the bladeplatform 36 to the blade tip 38. The leading edge section 50 includes apredetermined first width W1 such that the leading edge section 50encompasses nicks 52 and tears that may occur along the leading edge ofthe airfoil 34. The airfoil 34 subject to a significant tensile stressfield due to centrifugal forces generated by the fan blade 8 rotatingduring engine operation. The airfoil 34 is also subject to vibrationsgenerated during engine operation and the nicks 52 and tears operate ashigh cycle fatigue stress risers producing additional stressconcentrations around them.

To counter fatigue failure of portions of the blade along possible cracklines that can develop and emanate from the nicks and tears at least oneand preferably both of the pressure side 46 and the suction side 48 havea laser shock peened surfaces 54 and a pre-stressed region 56 havingdeep compressive residual stresses imparted by laser shock peening (LSP)extending into the airfoil 34 from the laser shock peened surfaces asseen in FIG. 3. Preferably, the pre-stressed regions 56 are coextensivewith the leading edge section 50 in the chordwise direction to the fullextent of width W1 and are deep enough into the airfoil 34 to coalescefor at least a part of the width W1. The pre-stressed regions 56 areshown coextensive with the leading edge section 50 in the radialdirection along the leading edge LE but may be shorter. The entire lasershock peened surfaces 54 is formed by overlapping laser shocked peenedcircular spots 58.

FIG. 2A illustrates the invention for a partial leading edge lengthpre-stressed regions 56 extending over a laser shock peened surfacelength Li of the leading edge LE that is generally centered about apredetermined nodal line 59 where it intercepts the leading edge LE.Preferably, the nodal line 59 is one of a dominant failure mode due tovibratory stress. This stress may be due to excitations of the blade inbending and torsional flexure modes. The dominant failure mode may notalways be the maximum stress mode but rather a lower stress mode orcombination of modes that exist for longer durations over the engine'smission. By way of example the predetermined nodal line 59 illustratedin FIG. 2A is due to a first flex mode. A nick 52 located in this areaof the leading edge LE has the greatest potential for failing the bladeunder resonance in this mode. Further by way of example, the laser shockpeened surface length Li of the partial leading edge length pre-stressedregions 56 may extend along the leading edge LE about 20% of the fanblade length from the tip 38 to the platform 36.

The present invention includes a fan blade construction with only thetrailing edge TE having laser shock peened surfaces 54 on a trailingedge section 70 having a second width W2 and along the trailing edge TE.The associated pre-stressed regions 56 having deep compressive residualstresses imparted by laser shock peening (LSP) extends into the airfoil34 from the laser shock peened surfaces 54 on the trailing edge section70. At least one and preferably both of the pressure side 46 and thesuction side 48 have a laser shock peened surfaces 54 and a pre-stressedregion 56 having deep compressive residual stresses imparted by lasershock peening (LSP) extending into the airfoil 34 from the laser shockpeened surfaces on a trailing edge section along the trailing edge TE.Preferably, the compressive pre-stressed regions 56 are coextensive withthe leading edge section 50 in the chordwise direction to the fullextent of width W2 and are deep enough into the airfoil 34 to coalescefor at least a part of the width W2. The compressive pre-stressedregions 56 are shown coextensive with the leading edge section 50 in theradial direction along the trailing edge TE but may be shorter,extending from the tip 38 a portion of the way along the trailing edgeTE towards the platform 36.

Referring to FIGS. 6 and 7, the laser beam shock induced deepcompressive residual stresses in the compressive pre-stressed regions 56are generally about 50-150 KPSI (Kilo Pounds per Square Inch) extendingfrom the laser shocked surfaces 54 to a depth of about 20-50 mils intolaser shock induced compressive residually stressed regions 56. Thelaser beam shock induced deep compressive residual stresses are producedby repetitively firing two high energy laser beams 2 each of which isdefocused ± a few mils with respect to the surfaces 54 on both sides ofthe leading edge LE which are covered with paint 55. The laser beamtypically has a peak power density on the order of magnitude of agigawatt/cm² and is fired through a curtain of flowing water that isflowed over the painted surface 54. The paint is ablated generatingplasma which results in shock waves on the surface of the material.These shock waves are re-directed towards the painted surface by thecurtain of flowing water to generate travelling shock waves (pressurewaves) in the material below the painted surface. The amplitude andquantity of these shockwave determine the depth and intensity ofcompressive stresses. The paint is used to protect the target surfaceand also to generate plasma. Illustrated in FIGS. 6 and 7 is anapparatus 1 which has the blade 8 mounted in a conventionally well knownrobotic arm 27 used to continuously move and position the blade toprovide laser shock peening “on the fly” in accordance with oneembodiment of the present invention. The laser shock peened surfaces 54on both the pressure and suction sides 46 and 48 respectively of theleading edge LE are painted with an ablative paint 55. Then the blade 8is continuously moved while continuously firing the stationary laserbeams 2 through a curtain of flowing water 21 on the surfaces 54 andforming overlapping laser shock peened circular spots 58. The curtain ofwater 21 is illustrated as being supplied by a conventional water nozzle23 at the end of a conventional water supply tube 19. The laser shockpeening apparatus 1 has a conventional generator 31 with an oscillator33 and a pre-amplifier 37 and a beam splitter 43 which feeds thepre-amplified laser beam into two beam optical transmission circuitseach having a first and second amplifier 39 and 41, respectively, andoptics 35 which include optical elements that transmit and focus thelaser beam 2 on the laser shock peened surfaces 54. A controller 24 maybe used to modulate and control the laser beam apparatus 1 to fire thelaser beams 2 on the laser shock peened surfaces 54 in a controlledmanner. Ablated paint material is washed out by the curtain of flowingwater.

The laser may be fired sequentially “on the fly”, as illustrated in FIG.4, so that the laser shock peened surface 54 is laser shock peened withmore than one sequence of painting the surface and then continuouslymoving the blade while continuously firing the laser beam on the surfacesuch that adjacent laser shock peened circular spots are hit indifferent sequences. FIGS. 4 and 5 illustrates a pattern of lasershocked peened circular spots 58 (indicated by the circles) of four suchsequences S1 through S4. The S1 sequence is shown as full line circles,as opposed to dotted line circles of the other sequences, to illustratethe feature of having non adjacent laser shocked peened circular spots58 with their corresponding centers X along a row centerline 62. Thepattern of sequences entirely covers the laser shock peened surface 54.The laser shocked peened circular spots 58 have a diameter D in a row 64of overlapping laser shock peened circular spots. The pattern may be ofmultiple overlapping rows 64 of overlapping shock peened circular spotson the laser shock peened surfaces 54. A first overlap is betweenadjacent laser shock peened circular spots 58 in a given row and isgenerally defined by a first offset O1 between centers X of the adjacentlaser shock peened circular spots 58 and can vary from about 30%-50% ormore of the diameter D. A second overlap is between adjacent laser shockpeened circular spots 58 in adjacent rows and is generally defined by asecond offset O2 between adjacent row centerlines 62 and can vary fromabout 30%-50k of the diameter D depending on applications and thestrength or fluency of the laser beam. A third overlap in the form of alinear offset O3 between centers X of adjacent laser shock peenedcircular spots 58 in adjacent rows 64 and can vary from about 30%-50% ofthe diameter D depending on a particular application.

This method is designed so that only virgin or near virgin paint isablated away without any appreciable effect or damage on the surface ofthe airfoil. This is to prevent even minor blemishes or remelt due tothe laser which might otherwise cause unwanted aerodynamic effects onthe blade's operation. Several sequences may be required to cover theentire pattern and re-painting of the laser shock peened surfaces 54 isdone between each sequence of laser firings. The laser firing eachsequence has multiple laser firings or pulses with a period betweenfirings that is often referred to a “rep”. During the rep the part ismoved so that the next pulse occurs at the location of the next lasershocked peened circular spot 58. Preferably the part is movedcontinuously and timed to be at the appropriate location at the pulse orfiring of the laser beam. One or more repeats of each sequence may beused to hit each laser shocked peened circular spot 58 more than once.This may also allow for less laser power to be used in each firing orlaser pulse.

One example of the present invention is a fan blade 8 having an airfoilabout 11 inches long, a chord C about 3.5 inches, and laser shock peenedsurfaces 54 about 2 inches long along the leading edge LE. The lasershock peened surfaces 54 are about 0.5 inches wide (WI). A first row 64of laser shocked peened circular spots 58 nearest the leading edge LEextends beyond the leading edge by about 20% of the laser spot diameterD which is about 0.27″ thus imparting deep compressive residual stressesin the pre-stressed region 56 below the laser shock peened surfaces 54which extend about 0.54 inches from the leading edge. Four sequences ofcontinuous laser firings and blade movement are used. The firingsbetween reps of the laser are done on spots 58 which lie on unablatedpainted surfaces which requires a repaint between each of the sequences.Each spot 58 is hit three times and therefore three sets of foursequences are used for a total of twelve paint and repaints of the lasershock peened surface 54.

Illustrated in FIG. 5 is an alternative embodiment of a laser shockpeened process in accordance with the present invention. The process maybe used to laser shock peen the entire leading edge as illustrated inFIG. 2 or a portion of the leading edge of the fan blade a shown in FIG.2A using five rows of laser shock peened spots and covering the entirearea of laser shock peened surfaces 54 in four sequences designated S1,S2, S3 and S4. The laser shock peening process starts with the firstsequence where every four spots is laser shock peened on sequence 1while the blade is continuously moved and the laser beam is continuouslyfired or pulsed and the laser. The part is timed to move betweenadjacent laser shock peened spots in the given sequence such as S1. Thetiming coincides with the rep between the pulses of the continuous laserfiring on the blade. All five rows of the overlapping laser shockedpeened circular spots 58 contain spots of each sequence spaced apart adistance so that other laser shock peened circular spots of the samesequence don't effect the paint around it. Sequence 1, preceded by afirst painting, is shown by the complete or full circles in the FIG. 4while the other laser shock peened spots such as in sequence S2, S3 andS4 are illustrated as dotted line, single dashed line, and double dashedline circles respectively. Before the next sequence, such as betweensequence S1 and sequence S2, the entire area of the laser shock peenedsurfaces 54 to be laser shock peened is repainted. This procedure ofrepainting avoids any of the bare metal of the laser shock peenedsurface from being hit directly with the laser beam. For an areacoverage of five rows with the spacing between rows and between adjacentspots of about 30%, it is found that one paint and three repaints willbe necessary so that the part is actually painted four times in totalwhich is somewhat of a time consuming process. It has been founddesirable to laser shock peen a given part, such as a fan blade, withbetween two and five rows. It has also been found desirable to lasershock peen each spot 58 up to 3 or more times. If each spot 58 is hit 3times then 1 paint and 11 repaints is required for three sets ofsequences S1-S4 for a total of 12 paintings.

It has been found that the part can be laser shock peened without anypaint using on the fly laser shock peening which saves a considerableamount of time by not having to repaint. Furthermore, since it is oftendesired to laser shock peen each surface more than once and, inparticular, three times, it is possible to save twelve paintings of thesurface by laser shock peening without any paint at all. The laser shockpeening without paint may use a lower laser beam fluency or use the samepower level as with paint. The plasma that is created is made up of themetal alloy material of the blade or part itself. In such a case, aremelt will be left on the laser shock peened area after the laser shockpeening or pulsing part of the process is completed. This remelt willusually have to be removed in any one of many well known processes suchas by mechanical or chemical removing of the layer. The usefulness of anon-painted part with on the fly laser shock peening will depend on thethickness of the part and careful attention must particularly be givento thin airfoil leading and trailing edges. It has been found that twoto five rows of laser shock spots without painting is a very usefulnumber for fan and compressor blades. It should be noted that the plasmaand the metal alloy without paint rehardens and forms what is known as aremelt and therefore will require a removal in one of the well knownfashions.

The no paint “on the fly” laser shock peening process of the presentinvention forms the above mentioned plasma with metal alloy on thesurface of the airfoil and a single continuous sequence of overlappingspots 58 may be used, as illustrated in FIG. 6, for as many rows asdesired. After the laser shock peening process is done the laser shockpeened surface may be removed to a depth sufficient to remove the remeltthat forms on the surface that might interfere with the airfoilsoperation.

Referring more specifically to FIG. 3, the present invention includes afan blade 8 construction with either the leading edge LE or the trailingedge TE sections or both the leading edge LE and the trailing edge TEsections having laser shock peened surfaces 54 and associatedpre-stressed regions 56 with deep compressive residual stresses impartedby laser shock peening (LSP) as disclosed above. The laser shockedsurface and associated pre-stressed region on the trailing edge TEsection is constructed similarly to the leading edge LE section asdescribed above. Nicks on the leading edge LE tend to be larger thannicks on the trailing edge TE and therefore the first width W1 of theleading edge section 50 may be greater than the second width W2 of thetrailing edge section 70. By way of example W1 may be about 0.5 inchesand W2 may be about 0.25 inches.

While the preferred embodiment of the present invention has beendescribed fully in order to explain its principles, it is understoodthat various modifications or alterations may be made to the preferredembodiment without departing from the scope of the invention as setforth in the appended claims.

I claim:
 1. An on the fly method of laser shock peening a gas turbineengine part, said method comprising the following steps: continuouslymoving a metallic gas turbine engine part while continuously firing astationary laser beam, which repeatably pulses with relatively constantperiods between the pulses, on a portion of the part, using a laser beamwith sufficient power to vaporize material on a laser shock peenedsurface of the portion of the part with the pulses around laser beamspots formed by the laser beam on the surface to form a region havingdeep compressive residual stresses extending into the part from thelaser shock peened surface, flowing a curtain of water over the surfaceupon which the laser beam is firing while moving the part until thelaser shock peened surface is completely covered by the laser beam spotsat least once, wherein the part is moved and the laser beam is fired toproduce more than one row of overlapping circular laser beam spotshaving generally equally spaced apart linearly aligned center pointswherein adjacent rows of spots overlap, and wherein the laser shockpeened surface is laser shock peened using a set of sequences whereineach sequence comprises covering the surface with a material suitable togenerate a plasma which results in shock waves to form the region havingdeep compressive residual stresses and then continuously moving the partwhile continuously firing a stationary laser beam on the surface suchthat adjacent laser shock peened circular spots are hit in differentones of said sequences in said set.
 2. A method as claimed in claim 1wherein: the metallic gas turbine engine part is a gas turbine engineblade having a leading edge and a trailing edge, the portion of the partis one of the edges, and the laser shock peened surface is on at least apart of one of the edges extending radially along at least a portion ofsaid one of the edges.
 3. A method as claimed in claim 2 furthercomprising simultaneously laser shock peening two laser shock peenedsurfaces each of which is on one of two sides of the part bycontinuously moving the part while continuously firing using twostationary laser beams which repeatably pulse between relativelyconstant periods, on the portion of the blade, using the laser beamswith sufficient power to vaporize material on the two surfaces of theportion of the blade with the pulses around laser beam spots formed bythe laser beam on the surfaces to form regions having deep compressiveresidual stresses extending into the blade from the laser shock peenedsurfaces, and flowing a curtain of water over the surfaces upon whichthe laser beam is firing while moving the blade until the laser shockpeened surfaces are completely covered by laser beam spots at leastonce.
 4. A method as claimed in claim 2 wherein the part is movedlinearly to produce at least one row of overlapping circular laser beamspots having generally equally spaced apart linearly aligned centerpoints.
 5. A method as claimed in claim 4 wherein the part is moved andthe laser beam is fired to produce more than one row of overlappingcircular laser beam spots having generally equally spaced apart linearlyaligned center points wherein adjacent rows of spots overlap.
 6. Amethod as claimed in claim 4 wherein the laser beam is fired and thepart moved so that the center points of adjacent spots in adjacent rowsare offset from each other a generally equal amount in a direction alonga line on which the center points are linearly aligned.
 7. A method asclaimed in claim 5 further comprising simultaneously laser shock peeningtwo laser shock peened surfaces each of which is on one of two sides ofthe part by continuously moving the part while continuously firing usingtwo stationary laser beams which repeatably pulse between relativelyconstant periods, on the portion of the blade, using the laser beamswith sufficient power to vaporize material on the two surfaces of theportion of the blade with the pulses around laser beam spots formed bythe laser beam on the surfaces to form regions having deep compressiveresidual stresses extending into the blade from the laser shock peenedsurfaces, and flowing a curtain of water over the surfaces upon whichthe laser beam is firing while moving the blade until the laser shockpeened surfaces are completely covered by laser beam spots at leastonce.
 8. A method as claimed in claim 2 wherein a first row of the lasershocked peened circular spots nearest the edge extends beyond said oneof the edges.
 9. A method as claimed in claim 8 wherein the part ismoved linearly to produce at least one row of overlapping circular laserbeam spots having generally equally spaced apart linearly aligned centerpoints.
 10. A method an claimed in claim 9 wherein the part is moved andthe laser beam is fired to produce more than one row of overlappingcircular laser beam spots having generally equally spaced apart linearlyaligned center points wherein adjacent rows of spots overlap.
 11. Amethod as claimed in claim 10 wherein the laser beam in fired and thepart moved so that the center points of adjacent spots in adjacent rowsare offset from each other a generally equal amount in a direction alonga line on which the center points are linearly aligned.
 12. An on thefly method of laser shock peening a gas turbine engine part, said methodcomprising the following steps: continuously moving a metallic gasturbine engine part while continuously firing a stationary laser beam,which repeatably pulses with relatively constant periods between thepulses, on a portion of the part, using a laser beam with sufficientpower to vaporize material on a laser shock peened surface of theportion of the part with the pulses around laser beam spots formed bythe laser beam on the surface to form a region having deep compressiveresidual stresses extending into the part from the laser shock peenedsurface, flowing a curtain of water over the surface upon which thelaser beam is firing while moving the part until the laser shock peenedsurface is completely covered by the laser beam spots at least once,wherein the part is moved and the laser beam is fired to produce morethan one row of overlapping circular laser beam spots having generallyequally spaced apart linearly aligned center points wherein adjacentrows of spots overlap, wherein the laser shock peened surface is lasershock peened using a set of sequences wherein each sequence comprisescontinuously moving the part while continuously firing a stationarylaser beam on the surface to form every other spot in each of saidadjacent rows such that adjacent laser shock peened circular spots inthe each of said adjacent row are hit in different ones of saidsequences wherein the part is moved and the laser beam is fired toproduce more that one row of overlapping circular laser beam spots morethan one row of overlapping circular laser beam spots having generallyequally spaced apart linearly aligned center points wherein adjacentrows of spots overlap, and the laser shock peened surface is laser shockpeened using a set of sequences wherein each sequence comprises paintingthe surface such that the material on the surface is a paint suitable togenerate a plasma which results in shock waves to form the region havingdeep compressive residual stresses and then continuously moving the partwhile continuously firing a stationary laser beam on the surface suchthat adjacent laser shock peened circular spots are hit in differentones of said sequences in said set.
 13. A method as claimed in claim 12wherein the laser beam is fired and the part moved so that the centerpoints of adjacent spots in adjacent rows are offset from each other agenerally equal amount in a direction along a line on which the centerpoints are linearly aligned.
 14. A method as claimed in claim 13 whereineach spot is hit more than one time using more than one set of saidsequences.